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The Role of Nuclear Power

and Nuclear Propulsion in the

Peaceful Exploration of Space

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THE ROLE OF NUCLEAR POWER AND NUCLEAR PROPULSION IN THE PEACEFUL EXPLORATION

OF SPACE

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The following States are Members of the International Atomic Energy Agency:

The Agency’s Statute was approved on 23 October 1956 by the Conference on the Statute of the IAEA held at United Nations Headquarters, New York; it entered into force on 29 July 1957.

The Headquarters of the Agency are situated in Vienna. Its principal objective is “to accelerate and enlarge the contribution of atomic energy to peace, health and prosperity throughout the world’’.

AFGHANISTAN ALBANIA ALGERIA ANGOLA ARGENTINA ARMENIA AUSTRALIA AUSTRIA AZERBAIJAN BANGLADESH BELARUS BELGIUM BENIN BOLIVIA

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THAILAND

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UNITED STATES OF AMERICA URUGUAY

UZBEKISTAN VENEZUELA VIETNAM YEMEN ZAMBIA ZIMBABWE

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THE ROLE OF NUCLEAR POWER AND NUCLEAR PROPULSION IN THE PEACEFUL EXPLORATION

OF SPACE

INTERNATIONAL ATOMIC ENERGY AGENCY VIENNA, 2005

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IAEA Library Cataloguing in Publication Data

The role of nuclear power and nuclear propulsion in the peaceful exploration of space. — Vienna : International Atomic Energy Agency, 2005.

p. ; 24 cm.

STI/PUB/1197 ISBN 92–0–107404–2

Includes bibliographical references.

1. Nuclear propulsion. 2. Outer space — Exploration. I. International Atomic Energy Agency.

IAEAL 05–00413

COPYRIGHT NOTICE

All IAEA scientific and technical publications are protected by the terms of the Universal Copyright Convention as adopted in 1952 (Berne) and as revised in 1972 (Paris). The copyright has since been extended by the World Intellectual Property Organization (Geneva) to include electronic and virtual intellectual property. Permission to use whole or parts of texts contained in IAEA publications in printed or electronic form must be obtained and is usually subject to royalty agreements. Proposals for non-commercial reproductions and translations are welcomed and will be considered on a case by case basis. Enquiries should be addressed by email to the Publishing Section, IAEA, at sales.publications@iaea.org or by post to:

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© IAEA, 2005 Printed by the IAEA in Austria

September 2005 STI/PUB/1197

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FOREWORD

This publication has been produced within the framework of the IAEA’s innovative reactor and fuel cycle technology development activities. It elucidates the role that peaceful space related nuclear power research and development could play in terrestrial innovative reactor and fuel cycle technology development initiatives. This review is a contribution to the Inter- Agency Meeting on Outer Space Activities, and reflects the stepped up efforts of the Scientific and Technical Subcommittee of the Committee on the Peaceful Uses of Outer Space to further strengthen cooperation between international organizations in space related activities.

Apart from fostering information exchange within the United Nations organizations, this publication aims at finding new potential fields for innovative reactor and fuel cycle technology development. In assessing the status and reviewing the role of nuclear power in the peaceful exploration of space, it also aims to initiate a discussion on the potential benefits of space related nuclear power technology research and development to the development of innovative terrestrial nuclear systems.

The IAEA expresses its appreciation to all those who contributed to this publication, in particular to J. Graham (ETCetera Assessments LLP, United States of America), V. Ionkin (Institute for Physics and Power Engineering, Russian Federation) and N.N. Ponomarev-Stepnoi (Kurchatov Institute, Russian Federation).

The IAEA officer responsible for this publication was A. Stanculescu of the Division of Nuclear Power.

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EDITORIAL NOTE

Although great care has been taken to maintain the accuracy of information contained in this publication, neither the IAEA nor its Member States assume any responsibility for consequences which may arise from its use.

The use of particular designations of countries or territories does not imply any judgement by the publisher, the IAEA, as to the legal status of such countries or territories, of their authorities and institutions or of the delimitation of their boundaries.

The mention of names of specific companies or products (whether or not indicated as registered) does not imply any intention to infringe proprietary rights, nor should it be construed as an endorsement or recommendation on the part of the IAEA.

The authors are responsible for having obtained the necessary permission for the IAEA to reproduce, translate or use material from sources already protected by copyrights.

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CONTENTS

1. INTRODUCTION . . . 1

2. REGIMES FOR THE USE OF NUCLEAR POWER IN SPACE EXPLORATION . . . 2

3. RADIOISOTOPE POWER DEVICES . . . 7

3.1. TEGs . . . 7

3.2. Thermionic converters . . . 13

3.3. Soviet/Russian TEG developments . . . 14

3.4. Safety . . . 17

3.5. Future applications . . . 17

4. REACTORS IN SPACE . . . 18

4.1. US experience . . . 18

4.1.1. Studies of on-board nuclear reactors . . . 20

4.2. Soviet/Russian experience . . . 23

4.2.1. Romashka NPS . . . 23

4.2.2. BUK NPS . . . 26

4.2.3. TOPAZ NPS . . . 29

4.2.4. Yenisey (TOPAZ-2) NPS . . . 31

5. NUCLEAR PROPULSION SYSTEMS . . . 34

5.1. US directions . . . 36

5.1.1. Safe affordable fission engine (SAFE) . . . 37

5.1.2. Heat pipe operated Mars exploration reactor (HOMER) . . . 38

5.2. Soviet/Russian directions . . . 40

5.2.1. IGR reactor . . . 41

5.2.2. IVG-1 experimental bench reactor . . . 42

5.2.3. IRGIT reactor . . . 45

5.3. Future applications . . . 47

6. SAFETY . . . 48

7. OTHER INTERNATIONAL SPACE PROGRAMMES . . . 50

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7.1. China . . . 50

7.2. France . . . 52

7.3. India . . . 53

7.4. Italy . . . 54

7.5. Japan . . . 55

8. THE FUTURE . . . 55

8.1. Nuclear energy for interplanetary missions . . . 55

8.2. International projects . . . 57

8.2.1. Mars Together . . . 57

8.2.2. TOPAZ-2 . . . 57

8.3. The road ahead . . . 60

8.4. USA: Future directions . . . 66

8.4.1. Variable specific impulse magneto-plasma rocket (VASIMR) . . . 68

8.4.2. Ion engines . . . 68

8.5. Russian Federation: Future directions . . . 69

8.5.1. Transport and energy module TEM . . . 70

8.5.2. Advanced thermionic NPS . . . 71

8.5.3. Advanced NPSs using the external energy conversion systems . . . 74

8.5.4. Advanced nuclear systems using lithium–niobium technology . . . 76

8.5.5. Gas core NTP . . . 77

8.5.6. Nuclear photon engines for deep space exploration . . . 79

8.6. New technologies through nuclear space systems engineering . . . 80

8.7. The value of space technology in terrestrial applications . . . 91

8.7.1. Research and development . . . 91

8.7.2. Products, equipment and materials . . . 92

8.8. Problems to be solved in space by the use of nuclear power . . 98

9. CONCLUSIONS . . . 101

10. LOOKING AHEAD . . . 104

APPENDIX I: EVALUATION OF MARS MISSIONS . . . 107

APPENDIX II: SPACECRAFT LAUNCHES INVOLVING RADIOISOTOPE SYSTEMS . . . 108

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APPENDIX III: SPACE ACHIEVEMENTS . . . 111

APPENDIX IV: CASSINI SPACECRAFT . . . 114

APPENDIX V: AMTEC . . . 115

APPENDIX VI: PROPULSION PERFORMANCE . . . 116

APPENDIX VII: THE ROVER PROGRAMME . . . 117

APPENDIX VIII: COMPARISON OF REACTOR SIZES. . . 118

APPENDIX IX: SOVIET NUCLEAR POWER SYSTEMS IN SPACE . . . 119

APPENDIX X: SPACE EXPLORATION AGENCIES AND CONTRACTORS . . . 121

APPENDIX XI: INTERNET REFERENCES . . . 125

ACKNOWLEDGEMENTS . . . 129

BIBLIOGRAPHY . . . 131

CONTRIBUTORS TO DRAFTING AND REVIEW . . . 133

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1. INTRODUCTION

It is more than 100 years since the Russian theoretician Konstantin Eduardovitch Ziolkovsky advocated the use of liquid fuel rockets for space exploration and almost 80 years since Robert Hutchings Goddard launched the first liquid fuel rocket at Auburn, Massachussetts, in the United States of America. Since then, rocket development has continued apace, largely through the endeavours of experimentalists such as Goddard, Willy Ley, Hermann Oberth, Wernher von Braun and other pioneers in both the German Society for Space Travel and the American Rocket Society.

Rocket research and development was given a major boost during the Second World War when the potential of the rocket engine to provide the motive force of a long range weapon delivery system was recognized. The result was the German V-2.

The trajectory of the V-2 took it to the edge of the upper atmosphere and the border of space; it can be regarded as being the first ‘space’ rocket.

Development of rocket technology gained momentum after the Second World War when both the USA and the former Soviet Union embarked on extensive programmes, culminating in the first satellite launch (Sputnik 1) in October 1957 and the first moon landing in July 1969.

Artificial satellites of necessity require their own power source. For many satellites this has taken the form of solar panels, whereby electricity is generated by the photovoltaic effect of sunlight on certain substrates, notably silicon and germanium. For satellites in earth orbit this a common method of generating power. However, the intensity of sunlight varies inversely with the cube of the distance from the sun, which means that a probe sent out to the neighbourhood of Jupiter would only receive a few per cent of the sunlight it would receive were it in earth orbit. In this case the solar panels would of necessity be so large as to be entirely impractical.

Such considerations lead to the development of alternative sources of power and heating which are completely independent of solar energy. One alternative involves the use of nuclear power systems (NPSs). These rely on the use of radioisotopes and are generally referred to as radioisotope thermo- electric generators (RTGs), thermoelectric generators (TEGs) and radio- isotope heater units (RHUs). These units have been employed on both US and Soviet/Russian spacecraft for more than 40 years. Examples of the use of these power sources on US probes over this period include Apollo, Viking, Pioneer, Voyager, Galileo, Ulysses and Cassini missions. None of these missions illustrate the utility of RTGs better than Pioneer 10, which was the first such probe to use power supplied solely from a radioisotope (238Pu).

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Pioneer 10 was launched from Cape Kennedy on 2 March 1972. It was the first interplanetary probe, successfully navigating the asteroid belt before making rendezvous with Jupiter and Saturn. The probe was equipped with an array of instruments for measuring such phenomena as the solar wind and the magnetic and radiation fields surrounding Jupiter. In fact, its discovery of the intense radiation fields surrounding Jupiter influenced the design of the Voyager and Galileo probes. Regarding Saturn, its instruments detected another ring and discovered two new satellites, as well as measuring the planet’s magnetic field.

What was apparently the spacecraft’s last signal was received on 22 January 2003 by the Jet Propulsion Laboratory’s Deep Space Network. An attempt to contact Pioneer 10 was made on 3 February 2003, but this failed. By this time the strength of the probe’s signal had degraded to such an extent that further communication was impossible. For much of its 30-year life Pioneer 10 had been in contact with earth and during this time had transmitted valuable information on Jupiter and Saturn and the outer reaches of the solar system. It is now some 13 billion kilometres from earth.

None of this would have been possible without the use of RTGs to provide electrical power and to maintain the components’ temperatures within their operational ranges.

The use of space NPSs is not restricted to the provision of thermal and electrical power. Considerable research has been devoted to the application of nuclear thermal propulsion (NTP). Such propulsion units will be capable of transferring significantly heavier payloads into earth orbit than is currently possible using conventional chemical propellants.

This publication reviews the development of NPSs and nuclear propulsion systems used in several national space programmes and details the units’ salient characteristics and other data (Appendices I–XI). It provides a history of the missions on which they were deployed and summarizes their advantages over other systems.

2. REGIMES FOR THE USE OF NUCLEAR POWER IN SPACE EXPLORATION

A space exploration mission requires power at many stages: for the initial launch of the space vehicle and for subsequent manoeuvering; for instrumen- tation and communication systems; for warming or cooling vital systems; for lighting; for experiments and many more uses, especially in manned missions.

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To date, chemical rocket thrusters have been used for launching. It would be tempting to believe that all power could be supplied by solar means since the sun is available and free. However, in many cases the mission may take place in the dark and large solar panels are not always suitable for a mission.

Figure 1 shows the regimes of possible space power applicability.

For short durations of up to a few hours, chemical fuels can provide energy of up to 60 000 kW, but for durations of a month use is limited to a kilowatt or less. Owing to the diffuse nature of solar power, it is not practicable to provide rapid surges of large amounts of energy. On the other hand, solar power is most efficient for power levels of some 10–50 kW for as long as it is needed.

Nuclear reactors can provide almost limitless power for almost any duration. However, they are not practicable for applications below 10 kW.

Radioisotopes are best used for continuous supply of low levels (up to 5 kW) of

NUCLEAR REACTORS

SOLAR RADIOISOTOPES

CHEMICAL 105

104

103

102

101

100

10-1

1 HOUR 1 DAY 1 MONTH 1 YEAR 10 YEARS

DURATION OF USE

ELECTRICAL POWER LEVEL (kW(e))

FIG. 1. Regimes of possible space power applicability. Source: Los Alamos National Laboratory.

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power or in combinations up to many times this value. For this reason, especially for long interplanetary missions, the use of radioisotopes for commu- nications and the powering of experiments is preferred.

Figure 2 shows that from any nuclear process, heat is emitted. This heat can then either be converted into electricity or it can be used directly to supply heating or cooling. The initial decay produces some decay products and the use of the thermal energy will provide some additional excess thermal energy to be rejected.

The nuclear process shown in Fig. 2 can either be a critical reactor or radioisotope fuel source such as plutonium oxide. In either case the heat can be converted to electricity either statically through thermoelectrics or a thermionic converter, or dynamically using a turbine generator in one of several heat cycles (Rankine, Stirling, Brayton). A classification of potential space applications of nuclear power is shown in Table 1. The nuclear workhorses for current space missions are the RTGs and the TEGs powered by radioisotopes in the Russian Federation that provide electricity through static (and therefore reliable) conversion at power levels of up to half a kilowatt, or more by combining modules.

Nuclear reactors have also been used in space, one by the USA in 1965 (SNAP-10A) successfully achieved orbit. The former Soviet Union routinely flew spacecraft powered by reactors: 34 had been launched prior to 1989 (see Appendix IX). A Soviet position paper stated that the investigation of outer space is “unthinkable without the use of nuclear power sources for thermal and electrical energy”. The USA agreed.

FIG. 2. Generic space NPS. Source: Los Alamos National Laboratory.

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The use of nuclear power in space is more than simply one of several power options. The choice of nuclear power can make deep space missions possible and much more efficient. For example, in a comparison between a typical chemical propulsion mission to Mars and one using nuclear propulsion, owing to the mass ratio efficiencies and the larger specific impulse1, the chemically powered mission took a planned total of 919 d and provided a stay of 454 d on the planet. By comparison, a nuclear powered mission was completed in 870 d while it provided 550 d on the planet (see Appendix I). The outward bound and return journeys TABLE 1. CLASSIFICATION OF NUCLEAR POWER TYPES BEING CONSIDERED FOR SPACE APPLICATION

(Source: Los Alamos National Laboratory)

NPS type Electrical power range

(module size)

Power conversion

RTG Up to 500 W(e) Static: thermoelectric

Radioisotope dynamic conversion generator

0.5–10 kW(e) Dynamic:

Brayton Organic Rankine Reactor systems:

Heat pipe Solid core Thermionics

10–1000 kW(e) Static:

Thermoelectric Thermionics Dynamic:

Brayton Rankine Stirling Reactor system:

Heat pipe Solid core

1–10 MW(e) Brayton

Rankine Stirling Reactor:

Solid core Pellet bed Fluidized bed Gaseous core

10–100 MW(e) Brayton (open loop) Stirling

Magnetic hydrodynamic

1 The specific impulse (a measure of rocket performance and measured in seconds) is the equivalent exhaust velocity divided by the acceleration due to gravity at sea level (9.8 m/s2).

The thrust (measured in Newtons or kilograms of force) is directly proportional to the specific impulse but the power needed to produce it is proportional to the square of the specific impulse.

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took 30% less time. In the ‘map’ of possibilities involving time and a variety of payloads, nuclear power wins most of the time.

The prospects for using NPSs in space are determined by their advantages over conventional solar photovoltaic and other power sources, including:

(a) Independence of the distance to the sun and orientation with respect to the sun.

(b) Compactness (a 10 MW solar array would require solar panels that cover an area of 68 000 m2 at the distance of Mars and 760 000 m2 at Jupiter and their size would render them impracticable).

(c) Better mass and size parameters when used on unmanned spacecraft, beginning with a power level of several tens of kilowatts.

(d) The capability of providing a power level two to three times greater with the NPS mass depending relatively weakly on the power improvement.

(e) Resistance to the earth’s radiation belts.

(f) The possibility of combining nuclear power with electrical thrusters to give the highest efficiency of specific impulse for thrust and of building power/propulsion systems on this basis to allow launch of payload masses two to three times greater than those possible with conventional chemical propellant orbital boosters. This can be achieved while supplying 50–100 kW of electrical power and more for onboard instrumentation over periods of 10 years or more.

The experience accumulated in developing space NPSs, electrical thrusters and NTPSs could, in the future, enable a number of quantitatively new exploration missions, such as round the clock all-weather radar surveil- lance and global telecommunication systems, including global systems for communication with moving objects. In the future, space NPSs and combined nuclear power/propulsion systems (NPPSs) with an electrical power level of several hundred kilowatts will enable such long term space missions as global environmental monitoring, production at facilities in space, supply of power for lunar and Martian missions, and others.

As a measure of power needs in space, a space shuttle consumes about 15 kW in orbit while the International Space Station (ISS) uses 75 kW.

Estimates for a Mars habitat range from 20–60 kW — not including propulsion.

A baseline Mars mission would require about 10 MW, but higher power means faster transportation. Thus, a 200 MW engine could theoretically reach Mars in 39 d. Such power is only available through advanced NPSs.

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3. RADIOISOTOPE POWER DEVICES

3.1. TEGS

The basic TEG is a simple device. It is based on an effect discovered by the German scientist Thomas Johann Seebeck in 1821. He found that when two dissimilar wires are connected at two junctions, and if one junction is kept hot while the other is cold, an electric current will flow in the circuit. Such a pair of junctions is called a thermocouple or thermoelectric couple.

The heat can be supplied from an isotope as in an RTG. The conversion of the heat is static. The device has no moving parts and is, therefore, very reliable and continues for as long as the radioisotope source produces a useful level of energy. The heat production is, of course, continually decaying but the radioi- sotope is custom selected to fit the intended use of the electricity and for its planned mission duration.

Figure 3 shows a hot shoe, through which radioisotopic heat is introduced, connecting the positive and negative legs. Some excess heat is rejected at the bottom and an electric current is generated.

A comparison between the predicted performance of a 150 W RTG over 12 years and its actual performance during that time is shown in Fig. 4.

RTGs have been used in 26 US and many Russian missions over the past forty years, as well as in the later French missions. They were originally installed in long term remote navigational and meteorological satellites, but RTGs have since been used in a variety of lunar and planetary missions. An RTG is a very versatile unit that can be custom designed for very specific appli- cations.

An example of an RTG SNAP is shown in Fig. 5. This is the SNAP-27 and Fig. 6 shows it being removed from the Lunar Excursion Module by astronaut Gordon Bean during the Apollo 12 mission to the moon in 1969. Five of these units were used to power experimental packages on the lunar surface. They were an ideal choice for long missions that required the supply of continuous power during both the lunar day and night. Each unit produced 63 W at the end of a year of service.

The US designed general purpose heat source (GPHS) comprises 238Pu fuel pellets encased in iridium shells (4 pellets each weighing 151 g) and 572 multiply redundant thermocouples made of silicon–germanium (see Fig. 7).

Each thermocouple can produce more than half a watt. However, for other missions, different fuel and different thermocouple materials can be used.

Moreover, RTGs can be used as modules of a total space auxiliary power

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system for both redundancy and for total power output. For the Galileo and Ulysses space missions, which had much higher power requirements than the lunar experiments, the GPHS–RTG was designed to provide 300 W of electrical power with a nominal fuel loading of 4.4 kW. It used 18 heat source modules.

Another design, the lightweight radioisotope heater unit (RHU), is shown in Fig. 8. These units provide temperature control for sensitive electrical components. Each includes a 2.68 g 238Pu dioxide fuel pellet producing 1 W, clad in platinum–rhodium and encased in a graphite capsule for protection in the event of an accident. The Galileo spacecraft had 120 of these lightweight units in addition to its GPHS. The Galileo spacecraft was launched on FIG. 3. Operating principle of the thermoelectric converter. Source: Rockwell International.

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TIME (Years)

POWER OUTPUT (W)

FIG. 4. Comparison between the predicted and actual performance of a 150 W RTG over a 12 year period.

HEAT REJECTION FINS

OUTER CASE

HERMETIC SEAL

THERMOPILE HOT FRAME

COLD FRAME

MOUNTING LEGS

HERMETIC SEAL

FUEL CAPSULE LATCH PLATE

RADIOISOTOPE FUEL CAPSULE

FIG. 5. The SNAP-27 system. Source: NASA.

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18 October 1989 and arrived at Jupiter on 7 December 1995. The mission was extended through 1999 to allow it to fly past Europa, Callisto and Io. These dates and the invaluable information fed back indicate the reliability of its on- board sources of thermal control and electricity generation.

Appendix II shows a listing of US and Russian spacecraft that have used RTGs (or radioisotope powered TEGs in the Russian Federation), the numbers of RTG systems and the reasons for those missions. Appendix III lists the successes of programmes supported by those power systems. These successes, with requirements for the supply of steady and reliable power for up FIG. 6. Removal of SNAP-27 from the Lunar Excursion Module by astronaut Gordon Bean during the Apollo 12 mission to the moon in 1969. Source: NASA.

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to 14 years in locations well beyond those which would allow the use of solar power, would not have been possible without RTGs.

The international2 Cassini mission to Jupiter and Saturn was equipped with 3 RTGs (see Appendix IV) which produced 885 W at the beginning of the mission and 633 W at the end. Cassini also had 82 small RHUs and there were 35 more on the Huygens probe, each producing 1 W of heat to keep nearby electronics warm. These contained a total of about 0.32 kg of 238Pu.

2 Partners: The National Aeronautics and Space Administration (NASA), the European Space Agency (ESA), the Italian Space Agency (Agenzia Spaziale Italiana – ASI) and there were a total of 17 countries involved.

FUEL CLAD GRAPHITE

AEROSHELL CAP

GRAPHITE IMPACT SHELL CBCF

DISC LOCK

MEMBER AEROSHELL

LOCK SCREW

FLOATING MEMBRANE

FUEL PELLET

GRAPHITE IMPACT SHELL

CBCF DISC

CBCF SLEEVE

FIG. 7. GPHS module assembly. Source: US Department of Energy/General Electric Co.

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For the future, a new advanced radioisotopic power system has been designed. It uses alkali metal thermal to electric conversion (AMTEC) technology to convert the heat produced by its plutonium heat source. The AMTEC cell (Appendix V) is made up of eight beta alumina solid electrolyte tubes connected in series. The end of the cell with the tubes is adjacent to the hot end of the heat source. At this end, liquid sodium is heated to a vapour state and the sodium atoms in the vapour are driven through the walls of the tubes and in so doing are stripped of an electron, thus creating positively charged sodium ions. The vapour is cooled and collected in a condenser at the cold end of the cell and the cycle is repeated as the sodium flows through the

‘artery’ towards the hot surface at the other end of the cell. The cell uses thermal shields in its upper section to reduce radiative bypass heat losses from the hot side components to the cold side condenser.

HEAT SHIELD END CAP

CAPSULE

INSULATOR MIDDLE TUBE

INSULATOR PLUG

HEAT SHIELD INSULATOR OUTER TUBE

INSULATOR INNER TUBE INSULATOR

PLUG

FIG. 8. Lightweight RHU.

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Leads are taken from the first and eighth tubes in series as the positive and negative leads for the cell. An explanatory cutaway diagram of the AMTEC system is shown in Appendix V.

This is an area of space research and development in which the latest ideas can be beneficial to various ongoing international innovative reactor technology research and development initiatives for terrestrial applications, particularly because older versions of these devices have already been used to provide power in remote situations, e.g. lighthouses and in the Arctic.

3.2. THERMIONIC CONVERTERS

Thermionic energy conversion is another method of transforming heat into electricity. It comprises a static device with a very hot emitter surface (typically at 1800 K) that ‘boils’ electrons across a small space (about 0.5 mm) to a cooler collector surface (typically at 1000 K). This action essentially creates an electrical engine with the electrons as a working fluid. There are factors preventing this engine from achieving its ultimate efficiency of the Carnot cycle. Among them are:

(a) Radiant heat transfer between the hot emitter and the cool collector;

(b) Space charge effects between the plates;

(c) Energy losses to the environment.

Much of the development programmes aim to overcome these difficulties.

The USA had a development programme targeting a 120 kW(e) power level with lifetimes of 10 000–20 000 h (limited by heat induced effects on the materials). The programme first tested converters in the reactor core (a thermionic reactor) but this programme was terminated in 1970. Work undertaken since has addressed usage separate from the reactor, resulting in the more efficient use of both the reactor and the thermionic converter.

Thermionic diodes include fuel that is firstly surrounded by the emitter surface and secondly surrounded by the collector surface with electrical connections at the bottom to connect to the next diode in series.

A thermionic reactor does not contain fuel rods releasing heat to a coolant but thermionic fuel elements (TFEs) directly generating electricity. As with a typical reactor, the fuel is critical and is controlled by rotating control drums. The temperature of the hot emitter plates is in turn dependent upon the reactor power level. These thermionic fuel rods are packaged in series, much like torch batteries. Designs have been produced showing these TFEs to be about 2.5 cm in diameter and up to 40 cm long.

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Gulf General Atomics tested the Thermionic Test Reactor employing carbide and oxide fuels between 1962 and 1973. Combinations of thermionic diodes and heat pipes (similar to the SP-100, see Section 4.1.1) provide challenging development problems but offer great potential.

This is another area of space research and development that can be beneficial to various ongoing international innovative reactor and fuel cycle research and development initiatives with terrestrial applications (see Section 8.7).

3.3. SOVIET/RUSSIAN TEG DEVELOPMENTS

In September 1965, TEGs (Orion-1 and Orion-2) based on 210Po were launched into a near earth orbit as components of the Cosmos-84 and Cosmos- 90 satellites. The choice of 210Po (with a specific thermal power of 141 W/g and half-life of 138 d) allowed for a compact design which incorporated silicon semiconductor converters with an electrical output of ~20 W. The service life was determined mainly by the half-life of 210Po and this could reach ~3000 h.

In the mid-1970s, research and development on a complex radionuclide (radioisotope) power system using 238Pu was initiated to support long term research of Mars. This power system, named VISIT, included an RTG with an electrical power output of about 40 W, the excess heat from which was transferred to a heat exchanger by pipes. However, VISIT power system development was limited to conducting terrestrial tests of its design and the fabrication of scale models, as well as thermal and electrical prototypes. During this era, key problems connected with the creation of radioisotope powered TEGs, or RTGs, for space were solved, namely:

(a) The production and processing of the 238Pu;

(b) The production of the cermet tablet fuel based on plutonium dioxide;

(c) The structural materials for the manufacture of the RHU (capsules with radionuclide), as well as their compatibility with the fuel composition over a wide temperature range;

(d) The RHU single elements’ design and the production process;

(e) Bench testing of the RHU.

In 1992, a thermoelectric mock-up (using an electric heater instead of the radionuclide) with an electrical output of 3.75 W at end-of-life was fabricated and tested (see Fig. 9). Its thermal power was 100 W. The tellurium, lead and germanium based alloy semiconductors in the thermoelectric battery were

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medium temperature ones with heat removed by thermal radiation from a ribbed casing. This RTG was proposed for use as a lander power source under ESA’s Leda lunar programme.

In the late 1990s, RTGs were used as the electrical power supply for the research probes to be landed on Mars as part of the Mars-96 international mission. The mission included long life small autonomous stations and probes (see Fig. 10). RTGs were needed to maintain equipment at design tempera- tures, to power equipment and to recharge a battery for communication with the orbiting spacecraft. Thus, 8.5 W 238Pu RHUs and a 200 mW(e) RTG named Angel were developed for the small autonomous station spacecraft. The Angel and RHU are unified products destined to be used both as self-contained units for equipment heating and as the initial heat source to provide the steady heat flow to a thermoelectric converter. The small autonomous station included two RHUs and two RTGs. The complex monoblock radionuclide power system had an electrical power output of about 400 mW and includes a storage battery, two RHUs with the thermal power of 8.5 W each and a thermoelectric converter. It was developed for the probes.

FIG. 9. Thermoelectric mock-up: (1) electric heater, (2) thermoelectric battery, (3) heat insulator, (4) casing, (5) rib. Source: Kurchatov Institute.

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The Angel’s cylindrical heat unit (d = 4 cm, h = 6 cm) includes a heat shield case and carbon heat insulation surrounding the radioisotope’s ampoule.

The ampoule contains about 17 g of 238Pu dioxide with an activity of 260 Ci. The ampoule has a dual structure. High corrosion resistant platinum–rhodium alloys are used for the inner ampoule casing that contains two ceramic tablets of 238Pu dioxide clad in iridium. The inner ampoule is hermetically welded and has a release mechanism for dealing with radiogenic helium resulting from the alpha decay of 238Pu. The load bearing outer cladding is fabricated from high strength tantalum–tungsten alloys. After hermetically welding, its surface is coated with multilayer refractory materials. Thus, the radionuclide heat source construction has a double containment in each capsule and the ampoule itself is additionally protected against outer thermal and impact attacks by heat resistant carbon materials.

The Angel radioisotope powered TEG or RTG was developed on the same basis. Semiconductor thermoelectric materials based on bismuth–

telluride alloys are used as a converter. The RTG generates an electrical power of about 200 mW at an operating voltage of 15 V at room temperature. The power for the small autonomous station equipment is supplied from the RTG through a nickel–cadmium buffer battery.

SMALL STATION

RTG(2) + RHU(2)

FIG. 10. Mars-96 project: small autonomous station layout. Source: Kurchatov Institute.

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Radioisotope powered TEGs (or RTGs) of milliwatt electrical power for space application, such as the Angel RTG and its modifications, are compact, reliable in operation and have low mass and size, which makes them convenient for probes. The RTG waste heat is enough to maintain design temperatures for equipment working in deep space environments.

Further, the RHU is also used for heating gas to warm the instrument module of the Lunokhod-1 and Lunokhod-2 stations. The heat sources’

thermal power is 900 W.

Research and development work on thermionic converters, together with the 238Pu based radionuclide heat sources, is ongoing with an emphasis being placed on improving energy conversion efficiencies from 8–10% to 10–14%.

Such generators using a thermionic converter with an electrical power of 75–150 W were proposed as the electrical power sources for a ‘rover’ vehicle under the Leda programme.

3.4. SAFETY

The safety of all RHU applications using 238Pu must take into account normal operation, emergency conditions during the launch and an uncon- trolled descent from the circular orbit. Maintaining the 238Pu capsule airtight under all possible conditions is the basis for ensuring radiation safety. The Angel RHUs were developed and designed to be consistent with the Principles Pertaining to the Use of Nuclear Power Sources in Space that were approved by the United Nations General Assembly in 1992.

3.5. FUTURE APPLICATIONS

Although the principles of thermoelectric and thermionic heating and power devices are very simple, continued development is producing more powerful and more compact designs. These designs have many terrestrial appli- cations and therefore the research and development in this area has synergies with innovative reactor technology development activities.

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4. REACTORS IN SPACE

As discussed in Section 2, while radioisotope powered systems are ideal for long term low power functions, nuclear reactors have the capability of producing almost unlimited power above a kilowatt for any length of mission.

The USA used one in 1965 in its SNAP-10A probe and the Soviet/Russian programme has routinely used them. Thirty-four nuclear powered Soviet spacecraft were launched between 1970 and 1989.

4.1. US EXPERIENCE

Early intentions were to use nuclear reactors both to power space launches and to supply onboard power needs. Considerable research on SNAP systems led to the launch of SNAP-10A (see Fig. 11) on an Atlas launch vehicle, salient details of which are as follows:

T/E CONVERTER RADIATORS THERMOELECTRIC PUMP

EXPANSION COMPENSATOR

SUPPORT LEG

REACTOR

SHIELD

STRUCTURE 8-RING STIFFENERS

LOWER SODIUM–POTASSIUM MANIFOLD

INSTRUMENTATION COMPARTMENT

FIG. 11. SNAP-10A system. Source: Atomics International.

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— Launch date: 3 April 1965 21:24 GMT. Launch site: Vandenberg. Launch complex: PALC2-4. Launch vehicle: SLV-3 Atlas/Agena D.

— Payload: SNAP-10A/Agena D. Mass: 440 kg. Class: Technology. Type: Ion engine. Agency: USAF/AEC. Perigee: 1270 km. Apogee: 1314 km. Incli- nation: 90.3 deg. Period: 111.4 min. COSPAR: 1965-027A. The spacecraft carried a SNAP-10A nuclear power source. The onboard nuclear reactor provided electrical power for a 1 kg force ion engine. The craft’s telemetry failed but the reactor itself operated well.

However, US Government policy changed and no more nuclear reactors were launched. Although the US sent only one nuclear reactor power source into space, the SNAP-10A, considerable work had already been done on two other reactors, SNAP-2 and SNAP-8. In the convention adopted by the USA, all RTG auxiliary power systems were denoted by odd numerals, while even numerals were reserved for nuclear reactors. All three reactors were similar but had different power levels (see Table 2).

TABLE 2. COMPARISON OF SNAP-2, SNAP-10A AND SNAP-8 REACTORS

Characteristic SNAP-2 SNAP-10A SNAP-8

Power (kW) 3 0.58 35

Design lifetime (a) 1 1 1

Reactor power (kW) 55 43 600

Reactor outlet (K) 920 833 975

Fuel and spectrum U–ZrH thermal U–ZrH thermal U–ZrH thermal

Coolant Na–K-78 Na–K-78 Na–K-78

Power conversion Rankine (Hg) Thermoelectric (Si–Ge)

Rankine (Hg)

Hot junction (K) 777

Cold junction (K) 610

Turbine inlet temperature (K)

895 950

Condenser temperature (K)

590 645

Unshielded weight (kg) 545 295 4545

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Unsolved design issues such as mercury corrosion and crud, and protection of stator windings, bearings and the pump made the reliability of SNAP-2 and SNAP-8 potentially poor and it was for this reason that the SNAP- 10A with its thermoelectric power conversion system was used in flight. The SNAP-10A power conversion system is shown in Fig. 12.

4.1.1. Studies of on-board nuclear reactors

There are many possible designs of nuclear reactors for use in space.

Advanced space mission requirements for high power levels (25 500 kW(e)) coupled with compact size and long lifetimes favour the use of the fast reactor spectrum with highly enriched fuel. One design for a liquid metal cooled space reactor, which is still a major contender for the future, is shown in Fig. 13.

This design is heavily dependent upon the designs of terrestrial liquid metal cooled fast reactors but is adapted for spacecraft in which the mission is power production rather than breeding or waste reduction.

The reactor needs to be small, restrained and not dependent upon gravity for its control, which would be normal on earth. Therefore, the design uses rotating beryllium control drums that have boron carbide absorber segments.

FLOW 14.5 gpm

'P 9 kPa

THERMAL POWER 1000 W AVE. RADIATOR TEMP. 583 K

ELECTRIC POWER 580 W

AVE. HOT JUNCTION TEMP. 777 K AVE. RADIATOR TEMP. 610 K EFFICIENCY, CONVERSION 1.43%

VOLTAGE 30 V

POWER CONVERSION SYSTEM

761 K

829 K 831 K

833 K

PUMP

POWER 43 kW

'P 72 K

REACTOR

FIG. 12. SNAP-10A power conversion system. Source: Atomics International.

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However, the materials technology and proof of that technology have been completed in the non-space liquid metal fast breeder reactor programme.

This example of a distributed liquid metal cooled reactor is merely one of many candidate systems that include several variants of solid core reactors (see Table 3).

For a mass density of 30 kg/kW(e) in a small reactor, outlet temperatures must be of the order of 1200–1500 K. This temperature objective defines both the form of the fuel and the coolant. For higher power requirements, in the 0.5–5.0 MW(e) range, fluidized bed and pellet bed reactors with gas cooling have been studied.

Apart from the nuclear reactor, a power plant includes shielding and a power conversion system, including converters and an excess heat rejection system.

In 1983, NASA, the US Department of Energy and several other agencies agreed to fund a joint programme, named SP-100 (see Fig. 14), to develop reactor system technology. This programme developed a power system that included a lithium cooled reactor coupled by heat pipes to thermoelectric converters. In this way the reactor could be used remotely from a manned spacecraft.

INSULATION (ZrO2) TIE-DOWN BARS (Mo–Re)

HEAT PIPE (Mo–Re/Li) FUEL LAYERS (UO2/Mo–Re)

DRUM BEARING REFLECTOR (Be)

SAFETY PLUG CAP (Be)

THERMAL INSULATION (MULTIFOIL)

CONTROL DRUMS (Be) ABSORBER SEGMENT (B4C) FUEL MODULES

RETAINER BANDS (Mo) DRIVE SHAFT REFLECTOR (BeO)

CORE CONTAINMENT (Mo–Re)

FIG. 13. Distributed cooled (liquid metal) space reactor. Source: Los Alamos National Laboratory.

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The thermoelectric panels contain converters, which use heat directly radiated to the hot shoes from the heat pipes. The hot shoes are made from molybdenum or lightweight carbon–carbon composites. The thermoelectric converters are distributed throughout the panels with the cold shoes serving as the heat rejection surfaces. The whole power system would be secured to the user spacecraft by a boom running through the centre of the heat pipes. These basic technologies and their evaluation are already contributing to the future US programme.

TABLE 3. EXAMPLES OF SOLID CORE NUCLEAR REACTOR SYSTEMS

Solid core type Variant 1 Variant 2 Variant 3

Integral heat transport reactor

Matrix fuel, gaseous (He) coolant

Pin fuel, Na–Li coolant

In-core cylindrical thermionics, Na–K coolant Distributed heat

transport reactor

Heat pipe wafer or coated particle fuel with heat pipes

Liquid metal wafer or coated particle fuel with electromagnetic pumps

In-core thermionics wafer or coated particle fuel with either electromagnetic pumps or heat pipes

FIG. 14. SP-100 nuclear power system (radioactively coupled system design). Source:

Los Alamos National Laboratory.

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4.2. SOVIET/RUSSIAN EXPERIENCE

The development of space NPSs with direct conversion of nuclear fission thermal power into electrical power started in the early to mid-1950s. The former Soviet Union’s first NPS with the direct (thermoelectric) conversion of nuclear fission heat into electricity was the terrestrial Romashka NPS. This NPS first operated in August 1964 and generated about 6100 kW·h of electrical energy over 15 000 h.

The BUK space thermoelectric NPS was created in the 1960s and has an electrical power output of about 3 kW. After the conclusion of tests in the early 1970s this NPS was put into operation in near earth orbits. From 1970 to 1988 there were 32 launches of these power systems (reactors) as a component of the Cosmos series of spacecraft (see Appendix IX).

The development of a space thermionic NPS was also undertaken in the early 1960s. The first successful power test of a terrestrial prototype of a thermionic reactor converter for the TOPAZ NPS was completed in 1970.

Further testing over the next two decades of thermionic reactor converter prototypes and TOPAZ NPS terrestrial prototypes made it possible to test the TOPAZ in flight. The first test flights of two TOPAZ NPS models as components of the Plasma-A spacecraft (Cosmos-1818 and Cosmos-1867) took place in February and July 1987. Along with the TOPAZ NPS, the development and testing of the Yenisey thermionic NPS was started in the second half of the 1960s. Since the power, mass and size parameters of this NPS were similar to those of the TOPAZ NPS, in the West it was referred to as TOPAZ-2.

On the basis of the experience gained with first generation of NPSs (BUK, TOPAZ, Yenisey), the development of the next generation thermionic systems (NPS-25, NPS-50 (Space Star) and NPS-100) was started in the mid- 1980s. The parameters of these systems meet the higher requirements in terms of the electrical power and lifetimes imposed by new space exploration targets as well as the latest safety requirements.

4.2.1. Romashka NPS

The main unit of the Romashka NPS is shown in Fig. 15. The Romashka NPS is a converter based on a fast reactor, in which the heat generated in the reactor core is conducted to a coaxially arranged TEG located on the radial reflector outer surface. The reactor core comprises a stack of 11 fuel elements;

the segmented fuel elements consisting of discs of uranium bicarbide with 90% enriched 235U. This is located within a graphite package, so built that a

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significant part of the heat from the core goes through the package body, thus reducing the temperature drop in the uranium bicarbide.

A radial beryllium reflector encloses the reactor. Graphite bushings are located between the core and the reflector to prevent reflector deformation at the high operating temperatures. The bushings are coated with silicon carbide and beryllium oxide to protect them from chemical interaction with beryllium.

The reactor end reflectors are also made of beryllium. The high temperature heat insulation made of foam graphite and multilayer graphitized fabric is FIG. 15. The Romashka NPS reactor converter layout: (1) radiator ribs, (2) thermo- electric elements, (3) control rod, (4) reactor vessel, (5) upper reflector, (6) reactor core, (7) radial reflector. Source: Kurchatov Institute.

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mounted on the reactor end walls to reduce the heat transfer. The combination allows the reactor to operate with a temperature of up to 2173 K in the centre of the core and between 1273 K and 1373 K on the reflector outer surfaces.

The reactor control system consists of four rods located in the radial reflector and in the lower end reflector. Two rods are used for automatic and manual control, whereas the other two, together with the movable end reflector, are used for reactor protection in case of emergency.

High temperature semiconductor grade silicon–germanium alloy (Si: 85% by mass, Ge: 15% by mass) is used in the TEGs. These are mounted inside the hermetically sealed steel vessel in four groups, each group having an independent power outlet. The cell comprises two thermopiles with the n- and p-conductivity joined together on the hot side by the molybdenum keyboard.

On the cold side, separate pairs are joined with each other in series by a copper bridge onto a common arm running the height of the generator.

To prevent the thermoelectric converters shorting, insulating plates of beryllium oxide are used on the hot and cold sides. To reduce heat loss, all clearances between the thermoelectric cell and hollows in the TEG structure are filled with a cotton-like quartz and a helium atmosphere. A total of 192 enamel coated fins, with an emissivity of at least 0.9, reject excess heat. Basic details of the Romashka NPS reactor converter characteristics are presented in Table 4.

TABLE 4. THE ROMASHKA NPS REACTOR CONVERTER CHARACTERISTICS

Characteristic Value

Reactor core diameter/height (by package) (mm) 241/351

Radial reflector outer diameter/height (mm) 483/553

Reactor load mass by uranium-235 (kg) 49

Total mass of the TEG (with the casing and radiator)

and reactor (without drives and control rods) (kg) 635 Reactor converter effective thermal power (without

taking into account the end wall spread of heat) (kW) 28.2 Reactor converter electrical output (at start-of-life) (W) 460–475 Electrical power reduction over a lifetime of 15 000 h 80%

Reactor converter terminal operating voltage (four groups

of thermoelectric converters connected in series) (V) 21

Number of thermoelectric converters in a TEG 3072

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The reactor converter generated an electric power of 460–475 W at a constant optimum outer load at start-of-life. By the end of testing (after approximately 15 000 h) the reactor converter electrical power had decreased to 80% of its initial power. This electrical power loss was mainly accounted for by an increase in the thermoelectric converter inner resistance owing to the diffusion processes operating at the graphite disc/silicon–germanium alloy interface resulting in the formation of a high resistance silicon carbide layer and, partly, to the failure of contacts on the hot side.

4.2.2. BUK NPS

The BUK NPS includes the reactor, the shielding and the conic/

cylindrical radiator located in series along the axis. The radiator comprises a system of ribbed pipes for coolant flow united by input and output collectors. It is located on the load bearing frame structure that is joined to the spacecraft.

The BUK NPS uses a small fast reactor which contains 37 fuel rods. The fuel is a highly enriched uranium–molybdenum alloy; the 235U load being about 30 kg. Longitudinally movable control rods are placed in the beryllium side reflector. A two-loop liquid metal heat removal system uses a eutectic alloy of sodium and potassium as coolant. The first loop’s coolant, heated to about 973 K, is supplied to the outer casing of the TEG. The TEG, the inner cavities of which are hermetically sealed and filled with inert gas, is located under the radiator, behind the reactor shielding. The second circuit coolant removes the excess heat to the radiator with the coolant maximum temperature at the radiator inlet being about 623 K. The TEG has two independent sections: one for the spacecraft users and an auxiliary one for the power-to-conduction type electromagnetic pumps used for both coolant loops. The BUK NPS layout is shown in Fig. 16.

There are two cascade thermoelectric converters in the TEG; the first made of a high temperature alloy, the second made of a medium temperature alloy. The nuclear reactor thermal power is limited to about 100 kW, from which the maximum electrical power generated is about 3 kW, representing an efficiency of 3%. The BUK NPS lifetime was extended in operation to 4400 h by which time the electrical parameters of the TEG had degraded. The conversion efficiency at 4400 h was about 90% of its start-of-life value.

Radiation safety is provided by two diverse systems:

(1) The basic safety system, the spacecraft component, relied on moving the spacecraft into a long term burial orbit, close to circular, at a height of more than 850 km. There, nuclear reactor fission products can decay safely to the level of natural radioactivity. The orbit change system is

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located in the spacecraft module and is mechanically joined to the nuclear power unit and separated from the spacecraft service module in low operational orbit. The orbit change system includes an off-line propulsion system with its own control systems and an off-line source of electrical power.

(2) The back-up emergency system provides for the dispersion of fuel, fission products and other materials with induced activity into the upper layers of the earth’s atmosphere. This system ejects the fuel element assembly either in the operational orbit or when the object, which includes the nuclear reactor, enters denser atmospheric layers. During the descent, aerodynamic heating, thermal destruction, melting, evaporation, oxidation, etc., are expected to disperse the fuel into particles that are sufficiently small as to pose no excess radiological hazard to the population or to the environment.

This backup system consists of control devices and actuating mechanisms that deform and destroy special flexible elements by the pressure of gases from cylinders. A diagram of the fuel element assembly ejection system from the reactor core is shown in Fig. 17.

The backup safety system was introduced into the BUK NPS after the failure of Cosmos-954 spacecraft’s change of orbit system. The spacecraft’s descent resulted in large radioactive fragments of wreckage being strewn in a line across northern Canada in 1978. Characteristics of the BUK NPS are shown in Table 5.

FIG. 16. The BUK NPS layout: (1) nuclear reactor, (2) liquid metal circuit pipeline, (3) reactor shielding, (4) liquid metal circuit expansion tanks, (5) radiator, (6) TEG, (7) load bearing frame structure. Source: Kurchatov Institute.

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TABLE 5. CHARACTERISTICS OF THE BUK NPS

Characteristic Value

Power (kW(e)) <3

Design lifetime (a) 1

Reactor power (kW) <100

Reactor outlet temperature (K) 973

Fuel and spectrum U–(90% enriched)–Mo, fast

Coolant Na–K eutectic

Power conversion Two cascade thermoelectric converter (Si–Ge)

Hot junction temperature (K) 623

Unshielded weight (kg) 900

FIG. 17. Diagram of the fuel element assembly ejection system for the BUK NPS: (1) tube plate, (2) fuel element assembly, (3) reactor vessel, (4) control rod, (5) reactor shielding, (6) side reflector, (7) actuating mechanism. Source: Kurchatov Institute.

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4.2.3. TOPAZ NPS

The TOPAZ NPS includes a thermionic reactor converter with a caesium vapour supply system and control drum drive unit, the reactor shielding, the radiator and the frame by which the system is joined to the spacecraft service module (Fig. 18). The automatic control system is placed in the hermetically sealed service module and connected to the related nuclear power unit systems by electrical service lines.

The core consists of 79 TFEs and four zirconium hydride moderator discs.

The TFEs and cooling channels are located in the moderator disc openings and form a system of five concentric rows. Five-cell TFEs with a three layer collector stack are used, with fission gas vented from emitter assemblies to the interelectrode gap. The TFEs are electrically connected so that they form the working section of 62 TFEs and the pump section of 17 TFEs. The pump section, where the TFEs are connected in parallel, is intended to energize the conduction type electromagnetic pump of the nuclear power unit’s heat removal system. The TFEs within this section are connected at both ends in the caesium vapour. The operating section terminal electrical output is about 6 kW at a voltage of ~32 V. The pumping section current is ~1200 A at a voltage of 1.1 V. Before the reactor converter is brought to the rated electrical power level, the electromagnetic pump is fed from the startup unit by a high current storage battery located behind the radiation shielding.

Twelve rotating cylinders (drums) located in the side reflector provide thermal power control, reactivity compensation and emergency shutdown.

FIG. 18. The TOPAZ NPS layout: (1) caesium vapour supply system and control drum drive unit, (2) thermionic reactor converter, (3) liquid metal circuit pipeline, (4) reactor shielding, (5) liquid metal circuit expansion tank, (6) radiator, (7) frame structure. Source:

Kurchatov Institute.

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These beryllium cylinders have sector cover plates of boron carbide and are divided into four groups of three cylinders. Each group is controlled by its own drive.

The caesium vapour supply system pumps vapour through the TFE interelectrode space at a flow rate of about 10 g/d. A pyrolized graphite trap absorbs used caesium and non-condensing impurities are ejected into space.

The NPS uses a lithium hydride reactor shield located in a hermetically sealed steel container with the inner load bearing elements.

The single circuit sodium–potassium heat removal system includes a radiator that has load carrying capacity and also serves as a structural member.

The radiator is designed as a system of D-shaped tubes placed hydraulically in parallel. The tubes are welded into the radiator O-ring collectors and supported by the load bearing elements. The tubes’ plane surface is soldered to a steel radiator which has a high emissivity coating. The area of the radiator is about 7 m2, which ensures rejection of at least 170 kW(th) at a coolant temperature of 880 K.

The automatic system controls the NPS to rated thermal and electrical power levels, maintains the working section current or coolant temperature at the rated level, maintains the voltage of ~28 V for the on-board equipment supply lines and can provide shutdown of the thermionic reactor converter.

A high speed controller that redistributes the direct current of the thermionic reactor converter section between the spacecraft and ballast loads controls the voltage. In the nominal operating mode, the rated current of the operating section and, consequently, its electrical power are sustained by correcting thermal power. As the efficiency degrades the coolant temperature rises to 880 K. After that, instead of maintaining the current, the automatic control system limits the coolant temperature. The thermal power then remains practically constant, whereas the operating section current will fall to values at which the onboard network voltage exceeds allowable limits, requiring NPS shutdown. NPS shutdown is also provided for specific emergency situations, as well as by radio command from earth.

The TOPAZ NPS generates approximately 6 kW at a start-of-life efficiency of about 5.5%. Its mass, including the nuclear power unit, the automatic control system and the coupling service lines, is about 1200 kg and it has a design lifetime of 4400 h. The nuclear power unit is 4.7 m long with a maximum diameter of 1.3 m.

In the period 1982–1984, two power tests of the TOPAZ NPS in combination with the automatic control system were performed in the automatic mode to prepare for flight tests. The first of these was performed with TFEs using emitter assemblies of single crystal molybdenum with a single crystal tungsten coating; the second with TFEs using single crystal

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molybdenum emitter assemblies. The first of the units was tested to ~4500 h and the second to ~7000 h. Results of testing fully corroborated the control algorithms in startup and operating modes, as well as agreeing with output parameters of the NPS, its subsystems and components and electrical power of no less than 5.6 kW for the prescribed lifetime of 4400 h.

In 1987, two experimental Plasma-A satellites (see Fig. 19) were launched with new generation TOPAZ reactors.

Safety was provided during the TOPAZ tests by placing the spacecraft in a circular operational orbit at a height in excess of 800 km. This orbit would provide sufficient decay time (350 a) for radioactive materials and fission products. The reactor was only made critical once the safe orbit had been attained. Details are summarized as follows:

— 2 February 1987 Cosmos-1818 Programme: Radar Ocean Reconnaissance Satellite RORSAT. Launch Site: Baikonur. Launch Vehicle: Tsyklon 2.

Mass: 3800 kg. Perigee: 790 km. Apogee: 810 km. Inclination: 65.0o.

— 10 July 1987 Cosmos-1867 Programme: RORSAT. Launch Site: Baikonur.

Mass: 3800 kg. Perigee: 797 km. Apogee: 813 km. Inclination: 65.0o. The NPS used with the first of the spacecraft (Cosmos-1818) operated for 142 d and the second one (Cosmos-1867) for 342 d. In both cases the NPS operation was terminated as planned when the caesium stock was exhausted.

The test programme objectives were fulfilled for both units. The flight test results confirmed the TOPAZ NPS output parameters and that operation in terrestrial conditions agreed with those in space. They attested to the stable operation of the reactor converter and its support systems in space flight and in the presence of operating plasma thrusters.

4.2.4. Yenisey (TOPAZ-2) NPS

Figure 20 shows a general view of the Yenisey NPS. All the equipment is packaged within a single unit referred to as the reactor or head unit, which has the shape of a truncated cone. The reactor is at the top, with the radiation

FIG. 19. Plasma-A satellite. Source: Kurchatov Institute.

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shielding located immediately underneath and all other equipment arranged in the shielding ‘shadow’.

The TOPAZ and Yenisey NPSs have similar structures and design arrangements. The principal difference between them is that the Yenisey thermionic reactor converter employs a single unit TFE; the emitter unit having an outer diameter of 19.6 mm and the collector pack an outer diameter of 23.7 mm (versus 10.0 mm and 14.6 mm, respectively, for TOPAZ). The main characteristics of the Yenisey NPS are presented in Table 6.

A single crystal molybdenum alloy with a single crystal tungsten 184 coating is used as the emitter material and the polycrystalline molybdenum alloy is used as the collector material. The emitter units have a central orifice through which the gaseous fission products are to be ejected into space. The TFEs are located in the thermionic reactor converter core tubes.

FIG. 20. General view of the Yenisey NPS. Source: Kurchatov Institute.

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The small clearance between the TFE and the tube is filled with helium.

In the reactor core there are 37 TFEs with the O-ring channels for their cooling located in orifices in zirconium hydride moderator discs. The operating section consists of 34 TFEs, the pumping section consisting of three. The electrical TABLE 6. MAIN CHARACTERISTICS OF THE YENISEY NPS

Description Value

Maximum electrical power at the reactor unit terminals supplied to

consumer (kW) 5.5

Current type Direct

Voltage (V) 27

Reactor thermal power (kW(th)) 135

Maximum coolant temperature at the reactor outlet (°C) 550

Maximum emitter temperature (°C) 1650

Lifetime corroborated by nuclear tests (a) 1.5

Reactor unit mass (kg) 1000

Dimensions of the reactor unit:

Length (mm) 3900

Maximum diameter (mm) 1400

Radiation situation over a plane of diameter 1.5 m at 6.5 m from the core centre:

Fluence of neutrons with energy >0.1 MeV (n/cm2 ) 5 × 1012

Gamma radiation exposure dose (R) 5 × 105

Core diameter (mm) 260

Core height (mm) 375

Number of TFEs in the core 37

Number of rotational control elements in the side reflector 12

Loading of uranium-235 in the core (kg) 25

Effective neutron multiplication factor (control elements out,

cold state) (keff) 1.005

Total reactivity temperature effect (Dk/k) 0.012

Worth of 12 control elements (Dk/k) 0.055

Peak to average power density:

Along to the core radius 1.1

Along to the core height 1.26

Lifetime ensured by the reactivity margin 3

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